Swept turbomachinery blade

ABSTRACT

A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade ( 12 ) has an airfoil ( 22 ) uniquely swept so that an endwall shock ( 64 ) of limited radial extent and a passage shock ( 66 ) are coincident and a working medium ( 48 ) flowing through interblade passages ( 50 ) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point ( 40 ) located at an inner transition radius r t -inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius r t-outer , radially inward of the airfoil tip ( 26 ), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.

CROSS-REFERENCE TO RELATED APPLICATIONS

This is a continuation of application Ser. No. 09/343,736, filed Jun.30, 1999 now U.S. Pat. No. Re. 38,040, seeking reissue of U.S. Pat. No.5,642,985, issued Jul. 1, 1997.

STATEMENT REGARDING GOVERNMENT RIGHTS

The government has certain rights to this invention under Department ofDefense Contract No. N00140-91-C-2793.

TECHNICAL FIELD

This invention relates to turbomachinery blades, and particularly toblades whose airfoils are swept to minimize the adverse effects ofsupersonic flow of a working medium over the airfoil surfaces.

BACKGROUND OF THE INVENTION

Gas turbine engines employ cascades of blades to exchange energy with acompressible working medium gas that flows axially through the engine.Each blade in the cascade has an attachment which engages a slot in arotatable hub so that the blades extend radially outward from the hub.Each blade has a radially extending airfoil, and each airfoil cooperateswith the airfoils of the neighboring blades to define a series ofinterblade flow passages through the cascade. The radially outerboundary of the flow passages is formed by a case which circumscribesthe airfoil tips. The radially inner boundary of the passages is formedby abutting platforms which extend circumferentially from each blade.

During engine operation the hub, and therefore the blades attachedthereto, rotate about a longitudinally extending rotational axis. Thevelocity of the working medium relative to the blades increases withincreasing radius. Accordingly, it is not uncommon for the airfoilleading edges to be swept forward or swept back to mitigate the adverseaerodynamic effects associated with the compressibility of the workingmedium at high velocities.

One disadvantage of a swept blade results from pressure waves whichextend along the span of each airfoil suction surface and reflect offthe surrounding case. Because the airfoil is swept, both the incidentwaves and the reflected waves are oblique to the case. The reflectedwaves interact with the incident waves and coalesce into a planaraerodynamic shock which extends across the interblade flow channelbetween neighboring airfoils. These “endwall shocks” extend radiallyinward a limited distance from the case. In addition, thecompressibility of the working medium causes a passage shock, which isunrelated to the above described endwall shock, to extend across thepassage from the leading edge of each blade to the suction surface ofthe adjacent blade. As a result, the working medium gas flowing into thechannels encounters multiple shocks and experiences unrecoverable lossesin velocity and total pressure, both of which degrade the engine'sefficiency. What is needed is a turbomachinery blade whose airfoil isswept to mitigate the effects of working medium compressibility whilealso avoiding the adverse influences of multiple shocks.

DISCLOSURE OF THE INVENTION

It is therefore an object of the invention to minimize the aerodynamiclosses and efficiency degradation associated with endwall shocks bylimiting the number of shocks in each interblade passage.

According to the invention, a blade for a blade cascade has an airfoilwhich is swept over at least a portion of its span, and the section ofthe airfoil radially coextensive with the endwall shock intercepts theendwall shock extending from the neighboring airfoil so that the endwallshock and the passage shock are coincident.

In one embodiment the axially forwardmost extremity of the airfoil'sleading edge defines an inner transition point located at an innertransition radius radially inward of the airfoil tip. An outertransition point is located at an outer transition radius radiallyintermediate the inner transition radius and the airfoil tip. The outertransition radius and the tip bound a blade tip region while the innerand outer transition radii bound an intermediate region. The leadingedge is swept at a first sweep angle in the intermediate region and isswept at a second sweep angle over at least a portion of the tip region.The first sweep angle is generally nondecreasing with increasing radiusand the second sweep angle is generally non-increasing with increasingradius.

The invention has the advantage of limiting the number of shocks in eachinterblade passage so that engine efficiency is maximized.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross sectional side elevation of the fan section of a gasturbine engine showing a swept back fan blade according to the presentinvention.

FIG. 2 is an enlarged view of the blade of FIG. 1 including analternative leading edge profile shown by dotted lines and a prior artblade shown in phantom.

FIG. 3 is a developed view taken along the line 3-3 of FIG. 2illustrating the tips of four blades of the present invention along withfour prior art blades shown in phantom.

FIG. 4 is a schematic perspective view of an airfoil fragmentillustrating the definition of sweep angle.

FIG. 5 is a developed view similar to FIG. 3 illustrating an alternativeembodiment of the invention and showing prior art blades in phantom.

FIG. 6 is a cross sectional side elevation of the fan section of a gasturbine engine showing a forward swept fan blade according to thepresent invention and showing a prior art fan blade in phantom.

FIG. 7 is a developed view taken along the line 7-7 of FIG. 6illustrating the tips of four blades of the present invention along withfour prior art blades shown in phantom.

BEST MODE FOR CARRYING OUT THE INVENTION

Referring to FIGS. 1-3, the forward end of a gas turbine engine includesa fan section 10 having a cascade of fan blades 12. Each blade has anattachment 14 for attaching the blade to a disk or hub 16 which isrotatable about a longitudinally extending rotational axis 18. Eachblade also has a circumferentially extending platform 20 radiallyoutward of the attachment. When installed in an engine, the platforms ofneighboring blades in the cascade abut each other to form the cascade'sinner flowpath boundary. An airfoil 22 extending radially outward fromeach platform has a root 24, a tip 26, a leading edge 28, a trailingedge 30, a pressure surface 32 and a suction surface 34. The axiallyforwardmost extremity of the leading edge defines an inner transitionpoint 40 at an inner transition radius r_(t)-inner, radially inward ofthe tip. The blade cascade is circumscribed by a case 42 which forms thecascade's outer flowpath boundary. The case includes a rubstrip 46 whichpartially abrades away in the event that a rotating blade contacts thecase during engine operation. A working medium fluid such as air 48 ispressurized as it flows axially through interblade passages 50 betweenneighboring airfoils.

The hub 16 is attached to a shaft 52. During engine operation, a turbine(not shown) rotates the shaft, and therefore the hub and the blades,about the axis 18 in direction R. Each blade, therefore, has a leadingneighbor which precedes it and a trailing neighbor which follows itduring rotation of the blades about the rotational axis.

The axial velocity V_(x) (FIG. 3) of the working medium is substantiallyconstant across the radius of the flowpath. However the linear velocityU of a rotating airfoil increases with increasing radius. Accordingly,the relative velocity V_(r) of the working medium at the airfoil leadingedge increases with increasing radius, and at high enough rotationalspeeds, the airfoil experiences supersonic working medium flowvelocities in the vicinity of its tip. Supersonic flow over an airfoil,while beneficial for maximizing the pressurization of the workingmedium, has the undesirable effect of reducing fan efficiency byintroducing losses in the working medium's velocity and total pressure.Therefore, it is typical to sweep the airfoil's leading edge over atleast a portion of the blade span so that the working medium velocitycomponent in the chordwise direction (perpendicular to the leading edge)is subsonic. Since the relative velocity V_(r) increases with increasingradius, the sweep angle typically increases with increasing radius aswell. As shown in FIG. 4, the sweep angle σ at any arbitrary radius isthe acute angle between a line 54 tangent to the leading edge 28 of theairfoil 22 and a plane 56 perpendicular to the relative velocity vectorV_(r). The sweep angle is measured in plane 58 which contains both therelative velocity vector and the tangent line and is perpendicular toplane 56. In conformance with this definition sweep angles σ₁ and σ₂,referred to hereinafter and illustrated in FIGS. 2, 3 and 6 are shown asprojections of the actual sweep angle onto the plane of theillustrations.

Sweeping the blade leading edge, while useful for minimizing the adverseeffects of supersonic working medium velocity, has the undesirable sideeffect of creating an endwall reflection shock. The flow of the workingmedium over the blade suction surface generates pressure waves 60 (shownonly in FIG. 1) which extend along the span of the blade and reflect offthe case. The reflected waves 62 and the incident waves 60 coalesce inthe vicinity of the case to form an endwall shock 64 across eachinterblade passage. The endwall shock extends radially inward a limiteddistance, d, from the case. As best seen in the prior art (phantom)illustration of FIG. 3, each endwall shock is also oblique to a plane 67perpendicular to the rotational axis so that the shock extends axiallyand circumferentially. In principle, an endwall shock can extend acrossmultiple interblade passages and affect the working medium enteringthose passages. In practice, expansion waves (as illustrated by therepresentative waves 68) propagate axially forward from each airfoil andweaken the endwall shock from the airfoil's leading neighbor so thateach endwall shock usually affects only the passage where the endwallshock originated. In addition, the supersonic character of the flowcauses passage shocks 66 to extend across the passages. The passageshocks, which are unrelated to endwall reflections, extend from theleading edge of each blade to the suction surface of the blade's leadingneighbor. Thus, the working medium is subjected to the aerodynamiclosses of multiple shocks with a corresponding degradation of engineefficiency.

The endwall shock can be eliminated by making the case wallperpendicular to the incident expansion waves so that the incident wavescoincide with their reflections. However other design considerations,such as constraints on the flowpath area and limitations on the caseconstruction, may make this option unattractive or unavailable. Incircumstances where the endwall shock cannot be eliminated, it isdesirable for the endwall shock to coincide with the passage shock sincethe aerodynamic penalty of coincident shocks is less than that ofmultiple individual shocks.

According to the present invention, coincidence of the endwall shock andthe passage shock is achieved by uniquely shaping the airfoil so thatthe airfoil intercepts the endwall shock extending from the airfoil'sleading neighbor and results in coincidence between the endwall shockand the passage shock.

A swept back airfoil according to the present invention has a leadingedge 28, a trailing edge 30, a root 24 and a tip 26 located at a tipradius r_(tip). An inner transition point 40 located at an innertransition radius r_(t)-inner is the axially forwardmost point on theleading edge. The leading edge of the airfoil is swept back by aradially varying first sweep angle σ₁ in an intermediate region 70 ofthe airfoil (in FIG. 2 plane 56 appears as the line defined by theplane's intersection with the plane of the illustration and in FIG. 3the tangent line 54 appears as the point where the tangent linepenetrates the plane of the Figure). The intermediate region 70 is theregion radially bounded by the inner transition radius r_(t)-inner andthe outer transition radius r_(t)-outer. The first sweep angle, as iscustomary in the art, is nondecreasing with increasing radius, i.e. thesweep angle increases, or at least does not decrease, with increasingradius.

The leading edge 28 of the airfoil is also swept back by a radiallyvarying second sweep angle σ₂ in a tip region 74 of the airfoil. The tipregion is radially bounded by the outer transition radius r_(t)-outerand a tip radius r_(tip). The second sweep angle is nonincreasing(decreases, or at least does not increase) with increasing radius. Thisis in sharp contrast to the prior art airfoil 22′ whose sweep angleincreases with increasing radius radially outward of the innertransition radius.

The beneficial effect of the invention is appreciated primarily byreference to FIG. 3 which compares the invention (and the associatedendwall and passage shocks) to a prior art blade (and its associatedshocks) shown in phantom. Referring first to the prior art illustrationin phantom, the endwall shock 64 originates as a result of the pressurewaves 60 (FIG. 1) extending along the suction surface of each blade.Each endwall shock is oblique to a plane 67 perpendicular to therotational axis, and extends across the interblade passage of origin.The passage shock 66 also extends across the flow passage from theleading edge of a blade to the suction surface of the blade's leadingneighbor. The working medium entering the passages is thereforeadversely influenced by multiple shocks. By contrast, the nonincreasingcharacter of the second sweep angle of a swept back airfoil 22 accordingto the invention causes a portion of the airfoil leading edge to be farenough forward (upstream) in the working medium flow that the section ofthe airfoil radially coextensive with the endwall shock extending fromthe airfoil's leading neighbor intercepts the endwall shock 64 (theunique sweep of the airfoil does not appreciably affect the location ororientation of the endwall shock; the phantom endwall shock associatedwith the prior art blade is illustrated slightly upstream of the endwallshock for the airfoil of the invention for illustrative clarity). Inaddition, the passage shock 66 (which remains attached to the airfoilleading edge and therefore is translated forward along with the leadingedge) is brought into coincidence with the endwall shock so that theworking medium does not encounter multiple shocks.

The embodiment of FIGS. 2 and 3 illustrates a blade whose leading edge,in comparison to the leading edge of a conventional blade, has beentranslated axially forward parallel to the rotational axis (thecorresponding translation of the trailing edge is an illustrativeconvenience—the location of the trailing edge is not embraced by theinvention). However the invention contemplates any blade whose airfoilintercepts the endwall shock to bring the passage shock into coincidencewith the endwall shock. For example, FIG. 5 illustrates an embodimentwhere a section of the tip region is displaced circumferentially(relative to the prior art blade) so that the blade intercepts theendwall shock 64 and brings it into coincidence with the passage shock66. As with the embodiment of FIG. 3, the displaced section extendsradially inward far enough to intercept the endwall shock over itsentire radial extent and brings it into coincidence with the passageshock 66. This embodiment functions as effectively as the embodiment ofFIG. 3 in terms of bringing the passage shock into coincidence with theendwall shock. However it suffers from the disadvantage that the airfoiltip is curled in the direction of rotation R. In the event that theblade tip contacts the rubstrip 46 during engine operation, the curledblade tip will gouge rather than abrade the rubstrip necessitating itsreplacement. Other alternative embodiments may also suffer from this orother disadvantages.

The invention's beneficial effects also apply to a blade having aforward swept airfoil. Referring to FIGS. 6 and 7, a forward sweptairfoil 122 according to the present invention has a leading edge 128, atrailing edge 130, a root 124 and a tip 126 located at a tip radiusr_(tip). An inner transition point 140 located at an inner transitionradius r_(t)-inner is the axially aftmost point on the leading edge. Theleading edge of the airfoil is swept forward by a radially varying firstsweep angle σ₁ in an intermediate region 70 of the airfoil. Theintermediate region is radially bounded by the inner transition radiusr_(t)-inner and the outer transition radius r_(t)-outer. The first sweepangle τ₁ is nondecreasing with increasing radius, i.e. the sweep angleincreases, or at least does not decrease, with increasing radius.

The leading edge 128 of the airfoil is also swept forward by a radiallyvarying second sweep angle σ₂ in a tip region 74 of the airfoil. The tipregion is radially bounded by the outer transition radius r_(t)-outerand the tip radius r_(tip). The second sweep angle is nonincreasing(decreases, or at least does not increase) with increasing radius. Thisis in sharp contrast to the prior art airfoil 122′ whose sweep angleincreases with increasing radius radially outward of the innertransition radius.

In the forward swept embodiment of the invention, as in the swept backembodiment, the nonincreasing sweep angle σ₂ in the tip region 74 causesthe endwall shock 64 to be coincident with the passage shock 66 forreducing the aerodynamic losses as discussed previously. This is incontrast to the prior art blade, shown in phantom where the endwallshock and the passage shock are distinct and therefore impose multipleaerodynamic losses on the working medium.

In the swept back embodiment of FIG. 2, the inner transition point isthe axially forwardmost point on the leading edge. The leading edge isswept back at radii greater than the inner transition radius. Thecharacter of the leading edge sweep inward of the inner transitionradius is not embraced by the invention. In the forward swept embodimentof FIG. 6, the inner transition point Is the axially aftmost point onthe leading edge. The leading edge is swept forward at radii greaterthan the inner transition radius. As with the swept back embodiment, thecharacter of the leading edge sweep inward of the inner transitionradius is not embraced by the invention. In both the forward swept andback swept embodiments, the inner transition point is illustrated asbeing radially outward of the airfoil root. However the invention alsocomprehends a blade whose inner transition point (axially forwardmostpoint for the swept back embodiment and axially aftmost point for theforward swept embodiment) is radially coincident with the leading edgeof the root. This is shown, for example, by the dotted leading edge 28″of FIG. 2.

The invention has been presented in the context of a fan blade for a gasturbine engine, however, the invention's applicability extends to anyturbomachinery airfoil wherein flow passages between neighboringairfoils are subjected to multiple shocks.

1. A turbomachinery blade for a turbine engine having a cascade ofblades rotatable about a rotational axis so that each blade in thecascade has a leading neighbor and a trailing neighbor, and each bladecooperates with its neighbors to define flow passages for a workingmedium gas, the blade cascade being circumscribed by a case and undersome operational conditions an endwall shock extends a limited distanceradially inward from the case and also extends axially andcircumferentially across the flow passages, and a passage shock alsoextends across the flow passages, the turbomachinery blade including anairfoil having a leading edge, a trailing edge, a root, a tip and aninner transition point located at an inner transition radius radiallyinward of the tip, the blade characterized in that at least a portion ofthe leading edge radially outward of the inner transition point is sweptand a section of the airfoil radially coextensive with the endwall shockextending from the leading neighbor intercepts the endwall shock so thatthe endwall shock and the passage shock are coincident.
 2. Aturbomachinery blade for a turbine engine having a cascade of bladesrotatable about a rotational axis so that each blade in the cascade hasa leading neighbor and a trailing neighbor, and each blade cooperateswith its neighbors to define flow passages for a working medium gas, theblade cascade being circumscribed by a case and under some operationalconditions an endwall shock extends a limited distance radially inwardfrom the case and also extends axially and circumferentially across theflow passages and a passage shock also extends across the flow passages,the turbomachinery blade including an airfoil having a leading edge, atrailing edge, a root, a tip located at a tip radius, an innertransition point located at an inner transition radius radially inwardof the tip, and an outer transition point at an outer transition radiusradially intermediate the inner transition radius and the tip radius,the blade having a tip region bounded by the outer transition radius andthe tip radius, and an intermediate region bounded by the innertransition radius and the outer transition radius, the bladecharacterized in that the leading edge is swept in the intermediateregion at a first sweep angle which is generally nondecreasing withincreasing radius, and the leading edge is swept over at least a portionof the tip region at a second sweep angle which is generallynonincreasing with increasing radius so that the section of the airfoilradially coextensive with the endwall shock extending from the leadingneighbor intercepts the endwall shock so that the endwall shock and thepassage shock are coincident.
 3. The turbomachinery blade of claim 1 or2 characterized in that the inner transition radius is coincident withthe root at the leading edge of the blade.
 4. A fan stage of a ductedfan gas turbine engine that is rotatable about an axis of rotation anddefines a downstream direction along the axis of rotation, comprising: afan casing that defines an inner duct wall having a fan rotor region; ahub disposed concentrically relative to the fan casing; a fan rotor thatincludes multiple swept fan blades, the swept fan blades being spacedapart around the hub, each of the multiple swept fan blades having: atip profile that corresponds to the inner duct wall of the fan casing; aleading edge that defines a variable sweep angle in a directionperpendicular to the axis of rotation, the leading edge including: aninner region adjacent the hub, the inner region defining a forward sweepangle; an intermediate region between the inner region and the fancasing, the intermediate region defining a rearward sweep angle; and anouter region between the intermediate region and the fan casing, theouter region being translated forward relative to a leading edge withthe same sweep angle as an outward boundary of the intermediate region.5. The fan stage according to claim 4, wherein the leading edge at aboundary between the intermediate region and the inner region extendsfurther upstream along the axis of rotation than the leading edge of theinner region.
 6. The fan stage according to claim 4, wherein each of themultiple swept fan blades includes a hub contacting surface that extendsfurther than the tip profile along the axis of rotation.
 7. A fan stageof a ducted fan gas turbine engine that is rotatable about an axis ofrotation and defines a downstream direction along the axis of rotation,comprising: a fan casing that defines an inner duct wall having a fanrotor region; a hub disposed concentrically relative to the fan casing;a fan rotor that includes multiple swept fan blades, the swept fanblades being spaced apart around the hub and being capable of rotatingat speeds providing supersonic working medium gas velocities over theblades to cause a shock in the gas adjacent the inner duct wall, each ofthe multiple swept fan blades having: a tip profile that corresponds tothe inner duct wall of the fan casing; a leading edge that defines avariable sweep angle in a direction perpendicular to the axis ofrotation, the leading edge including: an inner region adjacent the hub,the inner region defining a forward sweep angle; an intermediate regionbetween the inner region and the fan casing, the intermediate regiondefining a rearward sweep angle; and an outer region between theintermediate region and the fan casing, the outer region beingtranslated forward relative to a leading edge with the same sweep angleas an outward boundary of the intermediate region to provide a sweepangle that causes the blade to intercept the shock.